The present invention relates generally to gas turbine engines, and more particularly to hollow air cooled airfoils used in such engines.
A gas turbine engine includes a compressor that provides pressurized air to a combustor wherein the air is mixed with fuel and ignited for generating hot combustion gases. These gases flow downstream to one or more turbines that extract energy therefrom to power the compressor and provide useful work such as powering an aircraft in flight. In a turbofan engine, which typically includes a fan placed at the front of the core engine, a high pressure turbine powers the compressor of the core engine. A low pressure turbine is disposed downstream from the high pressure turbine for powering the fan. Each turbine stage commonly includes a stationary turbine nozzle followed in turn by a turbine rotor.
The turbine rotor comprises a row of rotor blades mounted to the perimeter of a rotor disk that rotates about the centerline axis of the engine. Each rotor blade typically includes a shank portion having a dovetail for mounting the blade to the rotor disk and an airfoil that extracts useful work from the hot gases exiting the combustor. A blade platform, formed at the junction of the airfoil and the shank portion, defines the radially inner boundary for the hot gas stream. The turbine nozzles are usually segmented around the circumference thereof to accommodate thermal expansion. Each nozzle segment has one or more nozzle vanes disposed between inner and outer bands for channeling the hot gas stream into the turbine rotor.
The high pressure turbine components are exposed to extremely high temperature combustion gases. Thus, the turbine blades and nozzle vanes typically employ internal cooling to keep their temperatures within certain design limits. The airfoil of a turbine rotor blade, for example, is ordinarily cooled by passing cooling air through an internal circuit. The cooling air normally enters through a passage in the blade""s root and exits through film cooling holes formed in the airfoil surface, thereby producing a thin layer or film of cooling air that protects the airfoil from the hot gases. Known cooling arrangements often include a plurality of openings in the trailing edge through which cooling air is discharged. These openings may take the form of holes, or of a pressure-side bleed slot arrangement, in which the airfoil pressure side wall stops short of the extreme trailing edge of the airfoil, creating an opening which is divided into individual bleed slots by a plurality of longitudinally extending lands incorporated into the airfoil casting. These slots perform the function of channeling a thin film of cooling air over the surface of the airfoil trailing edge. Airfoils having such a pressure-side bleed slot arrangement are known to be particularly useful for incorporating a thin trailing edge. In effect, the trailing edge thickness of the airfoil is equal to that of the suction side wall thickness alone. This is desirable in terms of aerodynamic efficiency.
Unfortunately, there are several problems associated with this geometry at the root of the airfoil. The trailing edge of most turbine blades is unsupported due to a large trailing edge overhang and is subject to mechanically and thermally induced stresses. The trailing edge blade root of an overhung blade is in compression considering mechanical stress only. In operation, there are significant thermal gradients from the trailing edge to adjacent regions of the blade, causing the trailing edge to be in compression considering thermal stress only. Additionally, in some designs the trailing edge root slot stops short of the blade fillet, resulting in a region below the root slot that is essentially uncooled which further exacerbates the radial and axial thermal gradients. The total stress of the trailing edge is the summation of these mechanical and thermal stresses, and can be at an undesirable level because both stress components are compressive. Because of these circumstances, the root slot is prone to thermal fatigue cracking.
Accordingly, there is a need for an airfoil having improved trailing edge root slot life.
The above-mentioned need is met by the present invention, which provides a hollow cooled turbine airfoil having pressure and suction side walls and a plurality of trailing edge cooling passages that feed cooling air bleed slots at the trailing edge. The airfoil trailing edge is selectively thickened in the root portion so as to allow shortened trailing edge slots, thereby improving trailing edge cooling and reducing mechanical stresses.
The present invention and its advantages over the prior art will become apparent upon reading the following detailed description and the appended claims with reference to the accompanying drawings.